Vibration control assembly

ABSTRACT

A vibration control assembly for an aircraft includes a housing operatively coupled to the aircraft. Also included is a cage disposed within an interior region of the housing, the cage rotatable within the housing about a first axis. Further included is a gyroscope wheel disposed within the cage and rotatable about a second axis other than the first axis, wherein a controllable moment is imposed on the aircraft upon rotation of the gyroscope wheel to counter vibratory moments produced by the vehicle. Yet further included is a control assembly at least partially surrounding the gyroscope wheel for controlling the controllable moment. The control assembly includes a structure having an inner surface, a track disposed along the inner surface, and an arm operatively coupled to the gyroscope wheel, the arm having an end disposed within the track, the gyroscope wheel angularly displaceable upon translation of the arm along the track

BACKGROUND OF THE DISCLOSURE

The embodiments herein generally relate to aircrafts and, moreparticularly, to a vibration control assembly for an aircraft andmethods of controlling aircraft vibration with one or more gyroscopeassemblies.

Helicopter rotors produce undesirable vibratory hub moments that causeunwanted fuselage vibration. The largest moments are in the pitching androlling moment directions. These moments produce an elliptical shape astime progresses. More generally, there are three moments which producean ellipsoidal shape. Typically, active vibration control techniques uselinear vibratory force actuators placed some distance apart in order tocreate a countering or anti-vibration moment. This approach undesirablyadds significant weight because the linear actuators rely upon linearlyoscillating a parasitic mass to generate load. However, the amplitudesof mass oscillation are limited due to space or other constraints,resulting in heavy designs that are deemed inefficient based on themoment produced relative to the weight. The reduction in payloadcapability of the aircraft is not desirably offset by the benefitsassociated with the counter-moment effects.

BRIEF DESCRIPTION OF THE DISCLOSURE

According to one embodiment, a vibration control assembly for anaircraft is provided. The assembly includes a housing operativelycoupled to the aircraft. Also included is a cage disposed within aninterior region of the housing, the cage rotatable within the housingabout a first axis. Further included is a gyroscope wheel disposedwithin the cage and rotatable about a second axis other than the firstaxis, wherein a controllable moment is imposed on the aircraft uponrotation of the gyroscope wheel to counter vibratory moments produced bythe vehicle. Yet further included is a control assembly at leastpartially surrounding the gyroscope wheel for controlling thecontrollable moment. The control assembly includes a structure having aninner surface, a track disposed along the inner surface, and an armoperatively coupled to the gyroscope wheel, the arm having an enddisposed within the track, the gyroscope wheel angularly displaceableupon translation of the arm along the track.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the structurecomprises a domed geometry.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the structure rotatessynchronously with the gyroscope wheel.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the track extendsalong the inner surface of the structure in a spiral path.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the gyroscope wheel isangularly displaceable over a 90 degree range.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the control assemblyfurther comprises a fin extending from an outer surface of the structureand a braking mechanism disposed proximate the fin and engageabletherewith to control a rotational speed of the structure.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the fin extends aroundat least a portion of a base of the structure.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the braking mechanismis an electric braking mechanism and is a regenerative brake that isconfigured to store energy during braking.

In addition to one or more of the features described above, or as analternative, further embodiments may include a motor operatively coupledto the cage with a motor shaft to rotate the cage and to controlprecession of the vibration control assembly.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the motor isoperatively coupled to the gyroscope wheel and drives rotation of thegyroscope wheel.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the motor is at leastpartially powered by energy generated by the regenerative brake of thebraking mechanism.

According to another embodiment, a method of controlling vibration on anaircraft is provided. The method includes rotating a cage about a cageaxis, the cage disposed within a housing. The method also includesrotating a gyroscope wheel about a gyroscope wheel axis that isnon-parallel to the cage axis, the gyroscope wheel disposed within thecage. The method further includes producing a moment on the aircraftupon rotating the gyroscope wheel, wherein the cage and gyroscope wheelpartially form a first vibration control assembly. The method yetfurther includes controlling the moment produced by controlling anangular orientation of the gyroscope wheel.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the angularorientation of the gyroscope wheel is controlled by translating an armoperatively coupled to the gyroscope wheel along a track formed along aninner surface of a structure that partially surrounds the gyroscopewheel.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the structure and thegyroscope wheel are rotated synchronously.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the rotational speedof the structure is controlled with a braking mechanism that applies abraking force to the structure.

In addition to one or more of the features described above, or as analternative, further embodiments may include that a power source isprovided power with energy generated by the braking force applied to thestructure.

In addition to one or more of the features described above, or as analternative, further embodiments may include controlling the momentproduced on the aircraft by varying a rotational speed of the gyroscopewheel

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter which is regarded as the invention is particularlypointed out and distinctly claimed in the claims at the conclusion ofthe specification. The foregoing and other features and advantages ofthe invention are apparent from the following detailed description takenin conjunction with the accompanying drawings in which:

FIG. 1 is a vibration control assembly according to an embodiment of theinvention;

FIG. 2 is the vibration control assembly according to another embodimentof the invention;

FIG. 3 is the vibration control assembly according to yet anotherembodiment of the invention;

FIG. 4 is the vibration control assembly according to yet anotherembodiment of the invention;

FIG. 5 illustrates instantaneous moment vectors of two vibration controlassemblies and their summed moment path to counter vibratory moments;

FIG. 6 is the vibration control assembly according to yet anotherembodiment;

FIG. 7 illustrates instantaneous moment vectors of four vibrationcontrol assemblies and their summed moment path to counter vibratorymoments;

FIG. 8 illustrates an aircraft using the vibration control assemblyaccording to an embodiment;

FIG. 9 is a side view of a control assembly for the vibration controlassembly, illustrating a first position of a gyroscope wheel;

FIG. 10 is a top view of the control assembly illustrating the gyroscopewheel in the first position;

FIG. 11 is a side view of the control assembly illustrating thegyroscope wheel in a second position; and

FIG. 12 is a top view of the control assembly illustrating the gyroscopewheel in the second position.

DETAILED DESCRIPTION OF THE DISCLOSURE

Referring to FIG. 1, illustrated is an active vibration control systemthat employs one or more vibratory control moment gyroscopes, alsoreferred to herein as a vibration control assembly 10. The vibrationcontrol assembly 10 is illustrated according to one embodiment inFIG. 1. It is contemplated that any structure, particularly vehicles,that inherently produces vibration may benefit from the counteractingvibratory forces of the embodiments described herein. One such structureis an aircraft, such as a helicopter that is subjected to vibration dueto moments generated by a rotor. The vibration control assembly 10counters the vibratory moments to reduce overall vibrations that thestructure (e.g., aircraft) is subjected to. In the case of a helicopter,the vibration control assembly 10 may be operatively coupled to alocation proximate the main transmission.

The vibration control assembly 10 includes a housing 12 that isoperatively coupled to the structure that is to undergo vibrationreduction. The housing may be operatively coupled to the structure inany suitable manner, including mounting with mechanical fasteners orwelded thereto. The housing 12 defines an interior region 14. A cage 16is disposed within the interior region 14. In the illustratedembodiment, the cage 16 is formed of an octagon cross-section, but it isto be appreciated that various alternative geometries may be employed.Irrespective of the geometry of the cage 16, a gyroscope wheel 18 isdisposed within the cage 16. The gyroscope wheel 18 includes a wheelsegment 20 and a gyroscope shaft 22, with the gyroscope shaft 22 coupledto the cage 16 at both ends. Gyroscope bearings 24 are included at bothends of the gyroscope shaft 22 to retain the gyroscope wheel 18.

The cage 16 is rotatable within the housing 12 about a first axis 26 andthe gyroscope wheel 18 is rotatable within the cage 16 about a secondaxis 28. The gyroscope bearings 24 are low friction bearings that allowthe gyroscope wheel 18 to rotate at a high rate (e.g., up to 20,000rpm), while also carrying a resulting gyroscopic moment that isproduced. The rotation of the gyroscope wheel 18 produces a moment thatpasses through the gyroscope bearings 24, through the cage 16 and intocage bearings 30 that are located at opposite ends of the cage 16proximate an interface between the cage 16 and the housing 12. Themoment is then passed to the housing 12 and subsequently to thestructure that the housing 12 is mounted to, thereby countering momentsproduced by the structure itself. The physics and dynamics of thecountering will be described in detail below.

A planetary gear arrangement 32 is employed to facilitate rotation ofthe gyroscope shaft 22 by operatively coupling the gyroscope shaft 22 toa driving source, thereby transmitting power to the planetary geararrangement 32. In the illustrated embodiment, the driving source is amotor 34 that may be an electric or hydraulic motor. Other drivingsources are contemplated. For example, a mechanical variable speed takeoff from the main transmission may be employed, such that the drivingsource is not only electrical or hydraulic. The motor 34 is operativelycoupled to a sun gear 36 with a motor shaft 38 that penetrates thehousing 12 and the cage 16. A planet gear shaft 40 is retained within agear retainer 42 within the cage 16 and includes a first planet gear 44and a second planet gear 46. The first planet gear 44 is disposed incontact with the sun gear 36 and the second planet gear 46 is disposedin contact with a drive gear 48 that is coupled to the gyroscope shaft22. This arrangement converts power from the motor 34 to rotationalmotion of the gyroscope shaft 22 about the second axis 28 and hence thegyroscope wheel 18. The motor speed varies dynamically between 0 andabout 20,000 rpm depending upon desired moment output which is monitoredand controlled by an outer-loop anti-vibration controller.

As noted above, the cage 16 is rotatable about a first axis 26. Rotationof the cage 16 is driven by a motor 50. The motor 50 includes a motorshaft 52 that is configured to penetrate the housing 12 and isoperatively coupled to the cage 16. The rotational speed and phase ofthe motor 50 is controllable. As will be appreciated, rotation of thecage 16 controls precession of the gyroscope wheel 18. As the cage 16 isrotated, the gyroscope wheel 18 and the planetary gear arrangement 32rotate with the cage 16, imparting precession of the gyroscope wheel 18.The combination of rotation of the gyroscope wheel 18 and the cage 16generates a gyroscopic moment 54. The magnitude of the gyroscopic moment54 is equal to the product of the precession speed, the gyroscope wheelspeed and the mass moment of inertia of the gyroscope wheel.

In the illustrated embodiment, the sun gear 36 and the motor shaft 38are coaxially aligned with each other and with the first axis 26 thatthe cage 16 rotates about. The first axis 26 coincides with the axis ofrotation of the motor shaft 52. As the planet gears 44, 46 rotate, theyorbit the sun gear 36. Motor torque is thus transmitted to the drivegear 48 that drives the gyroscope wheel 18. As torque is transmittedcontinuously to the planet gear 46, it transmits torque directly to thegyroscope shaft 22. Such an embodiment avoids the need for a rotatingmotor directly connected to the gyroscope wheel 18 to spin the gyroscopewheel 18, as well as obviates the need for slip rings to power such amotor.

Referring to FIG. 2, an alternative embodiment of the vibration controlassembly 10 is illustrated. In particular, the planetary geararrangement 32 is simplified. The motor 34 includes the motor shaft 38that extends through the housing 12 and the cage 16 into close proximitywith the gyroscope shaft 22 and has a gear 56 coupled thereto. The gear56 is connected to the drive gear 48, which directly rotates thegyroscope shaft 22.

Referring to FIG. 3, yet another embodiment of the vibration controlassembly 10 is illustrated. As shown, power may be transmitted to thegyroscope shaft 22 without a dedicated motor. Specifically, a gearboxshaft 60 that extends from an existing gearbox of the vehicle, such as amain rotor gearbox of a helicopter. The gearbox shaft 60 is powered byan existing power source and is simply disposed in contact with theplanetary gear arrangement 32 illustrated. The planetary geararrangement 32 of FIG. 3 includes two planet gear shafts 40 that areboth coupled to the sun gear 36. The embodiment described above andillustrated is merely exemplary, as one can appreciate that fewer ormore planet gear shafts 40 may be employed.

Referring to FIG. 4, another embodiment of the vibration controlassembly 10 is shown. In the illustrated embodiment, the motor 50 drivesrotation of both the cage 16 and the gyroscope wheel 18. To facilitatethe additional rotation of the gyroscope wheel 18, a slip ring 62 isdisposed between the motor 50 and the cage 16. Electrical power isconducted to a motor stator 64. Electric motor permanent magnets 68 areincluded about the outer perimeter of the wheel portion 20 of thegyroscope wheel 18 and at an inner portion of the motor stator 64 todrive rotation of the gyroscope wheel 18.

Referring now to FIG. 6, an embodiment of the vibration control assembly10 is shown with a single motor 70 driving rotation of the cage 16 andthe gyroscope wheel 18. In this embodiment, the sun gear 36 does notrotate which allows the gyroscope wheel 18 to be driven when thegyroscope wheel 18 is precessed, thereby avoiding the need for arotating motor directly connected to the gyroscope wheel 18 to spin thegyroscope wheel 18, as well as obviating the need for slip rings topower such a motor. The single motor 70 drives the gyroscope wheel 18 ata constant rotational speed in such an embodiment.

Irrespective of the precise configuration of the vibration controlassembly 10, it is to be appreciated that a single assembly may beincluded or a plurality of vibration control assemblies may be included.For example, two or four vibration control assemblies are provided insome embodiments.

Referring now to FIGS. 9-12, a control assembly 200 for any of theembodiments of the vibration control assembly 10 is illustrated in twopositions and therefore two different operating conditions of thevibration control assembly 10. As described herein, the control assembly200 varies the angle of one or more components of the vibration controlassembly 10 to control the amount of counter-vibration that thevibration control assembly 10 provides.

The control assembly 200 includes a structure 202 that surrounds all ora portion of the vibration control assembly 10. In the illustratedembodiment, the structure 202 is a dome-like structure, but it is to beunderstood that alternative geometries are contemplated. The structure202 includes an inner surface 204 that has a track 206 extendingtherealong. The track 206 may be built on the inner surface 204 or maybe defined as a recess within the inner surface 204. In the illustratedembodiment, the track 206 is oriented in a spiral configuration, but itis to be understood that the illustrated spiral configuration is merelyillustrative and is not limiting. In particular, it is contemplated thatthe track 206 may extend along the inner surface 204 in a linear manner,curvilinear manner or some other angular orientation.

The track 206 is sized to receive an end of an arm 208 therein. The arm208 is operatively coupled to the vibration control assembly 10. In theillustrated embodiment, the arm 208 is operatively coupled to thegyroscope wheel 18 (may also be referred to as a gimbal). The structure202 is configured to spin in a synchronous manner with the gyroscopewheel 18. As shown, in FIGS. 9 and 10, the gyroscope wheel 18 isdisposed in a first angular position (e.g., horizontal) when the arm 208is disposed in a top, central location of the track 206. The relativerotational rates of the gyroscope wheel 18 and the structure 202 maycause the arm 208 to translate along the track 206. Translation of thearm 208 along the track 206 causes a change in rotational angle of thegyroscope wheel 18. FIGS. 11 and 12 illustrate the gyroscope wheel 18 ina different angular position when compared to FIGS. 9 and 10.

The track 206 is formed of a material that has a coefficient of frictionthat is low enough to facilitate reliable translation of the arm 208 toavoid sticking. Alternatively, the track 206 may be coated with alow-friction material. In yet another alternative, the end of the arm208 may be formed of, or coated with, a low-friction material.Additionally, both the end of the arm 208 and the track 206 may includelow-friction material.

The structure 202 also includes a fin or the like 210 extending from anouter surface 212 of the structure 202. In the illustrated embodiment,the fin 210 is located proximate a base 214 of the structure 202. Thefin 210 may extend completely about the circumference of the base oronly along a portion thereof. A braking mechanism 216 is disposed inclose proximity with the fin 210, such as the surrounding configurationthat is illustrated. The braking mechanism 216 engages the fin 210 in acontrollable manner to exert a braking force on the structure 202 toslow rotation of the structure. The braking force allows an operator orcontroller to “dial in” the amount of counter vibration that thevibration control assembly 10 imposes based on the angular orientationof the gyroscope wheel 18. In some embodiments, the angular orientationof the gyroscope wheel 18 is moveable over a 90 degree range, such asfrom a horizontal orientation to a vertical orientation.

The braking mechanism 216 may be part of an electric, regenerativebraking system that is used to store energy that may be employed topower the vibration control assembly 10.

Referring to FIG. 5, the physics and dynamics of the resultant momentsimposed on the structure by two vibration control assemblies isillustrated at an instantaneous moment of time. Assuming that thegyroscope disks are rotating in the same direction, two assemblies(i.e., gyroscope wheels) that are forced to precess in oppositedirections at a desired vibration-suppression frequency can be used tocounter the undesirable moments produced by the structure (e.g.,vehicle). A vibration control assembly 80 is controlled to precess inthe clockwise direction at an angular speed of +Ω to produce a momentvector A. Vibration control assembly 82 is controlled to precess in thecounterclockwise direction at an angular speed of −Ω to produce a momentvector B. Each produces a circular rotating moment vector A, B, withboth vectors rotating or processing at the frequency of the undesiredvehicle vibration frequency. By various combinations of moment vectorsize (i.e., magnitude) and phase, any vibratory pitching or rollingmoment ellipse 84, at any “tilt” may be produced. Advantageously, theresulting ellipse (including a circle) may be adjusted to counter themoment produced by the structure to result in a zero net moment on thestructure, thereby producing extremely low vibration.

Referring to FIG. 7, a case where four vibration control assemblies areemployed is illustrated. Such embodiments may be beneficial where thegyroscope wheel 18 rotates at a nearly constant speed, such as in theembodiments described in FIGS. 2 and 6, for example. In theillustration, four assemblies are represented with wheel speed designedto have spin speeds that are fixed ratios of the gearbox in the exampleof FIG. 2. As a result, vibratory moments A, S, B and U are all the samesize. Two of the assemblies are made to precess in the same direction atthe same speed (−Ω), but the precession angular position of therespective gyroscope wheels are selected such that moment vector A andmoment vector S vectorially add so that the result is the desired momentT. Similarly, the other two assemblies are phased to produce net momentvector V through the vector addition of moment vector B and momentvector U. These two assemblies are made to precess in the oppositedirection from the other two assemblies, but at the same speed (i.e., inthe direction and precession speed +Ω. The two circular moments rotatingin opposite directions vectorially add to produce an elliptical momentpattern that may be controlled.

A single gyroscope of variable disk speed or two gyroscopes withconstant disk speed may also be advantageous on a vehicle that exhibitsa dominant, undesirable ambient vibration moment that is nearlycircular. This is an arrangement that is particularly advantageous whenit is desired to minimize vehicle weight at the cost of highervibrations.

Advantageously, effective anti-vibration is achieved at a reduced weightrequirement for assemblies employed to do so. The vibration controlassembly, or assemblies, efficiently generate large anti-rotationmoments by spinning the gyroscope wheel 18 faster, rather thanincreasing the travel of a linear-type actuator or the distance betweenlinear-type actuators.

FIG. 8 illustrates an aircraft using the vibratory control assemblyaccording to an embodiment. In particular, a rotary-wing aircraft 100having a dual, counter-rotating, coaxial rotor system 102 which rotatesabout a rotating main rotor shaft 14U, and a counter-rotating main rotorshaft 14L both about an axis of rotation A. The aircraft 100 includes anairframe F which supports the dual, counter rotating, coaxial rotorsystem 102 as well as an optional translational thrust system T whichprovides translational thrust during high speed forward flight generallyparallel to an aircraft longitudinal axis L. Although a particularaircraft configuration is illustrated in the disclosed embodiment, othercounter-rotating, coaxial rotor systems will also benefit from thepresent invention.

While the invention has been described in detail in connection with onlya limited number of embodiments, it should be readily understood thatthe invention is not limited to such disclosed embodiments. Rather, theinvention can be modified to incorporate any number of variations,alterations, substitutions or equivalent arrangements not heretoforedescribed, but which are commensurate with the spirit and scope of theinvention. By way of example, aspects can be used in conventional and/orcoaxial rotary aircraft, fixed wing aircraft, maritime applications,industrial machinery, automotive applications, or other applicationswhere vibrations need to be reduced. Additionally, while variousembodiments of the invention have been described, it is to be understoodthat aspects of the invention may include only some of the describedembodiments. Accordingly, the invention is not to be seen as limited bythe foregoing description, but is only limited by the scope of theappended claims.

1. A vibration control assembly for an aircraft comprising: a housingoperatively coupled to the aircraft; a cage disposed within an interiorregion of the housing, the cage rotatable within the housing about afirst axis; a gyroscope wheel disposed within the cage and rotatableabout a second axis other than the first axis, wherein a controllablemoment is imposed on the aircraft upon rotation of the gyroscope wheelto counter vibratory moments produced by the vehicle; and a controlassembly at least partially surrounding the gyroscope wheel forcontrolling the controllable moment, the control assembly comprising: astructure having an inner surface; a track disposed along the innersurface; and an arm operatively coupled to the gyroscope wheel, the armhaving an end disposed within the track, the gyroscope wheel angularlydisplaceable upon translation of the arm along the track.
 2. Thevibration control assembly of claim 1, wherein the structure comprises adomed geometry.
 3. The vibration control assembly of claim 1, whereinthe structure rotates synchronously with the gyroscope wheel.
 4. Thevibration control assembly of claim 1, wherein the track extends alongthe inner surface of the structure in a spiral path.
 5. The vibrationcontrol assembly of claim 1, wherein the gyroscope wheel is angularlydisplaceable over a 90 degree range.
 6. The vibration control assemblyof claim 1, wherein the control assembly further comprises: a finextending from an outer surface of the structure; and a brakingmechanism disposed proximate the fin and engageable therewith to controla rotational speed of the structure.
 7. The vibration control assemblyof claim 6, wherein the fin extends around at least a portion of a baseof the structure.
 8. The vibration control assembly of claim 6, whereinthe braking mechanism is an electric braking mechanism and is aregenerative brake that is configured to store energy during braking. 9.The vibration control assembly of claim 8, further comprising a motoroperatively coupled to the cage with a motor shaft to rotate the cageand to control precession of the vibration control assembly.
 10. Thevibration control assembly of claim 9, wherein the motor is operativelycoupled to the gyroscope wheel and drives rotation of the gyroscopewheel.
 11. The vibration control assembly of claim 9, wherein the motoris at least partially powered by energy generated by the regenerativebrake of the braking mechanism.
 12. A method of controlling vibration onan aircraft comprising: rotating a cage about a cage axis, the cagedisposed within a housing; rotating a gyroscope wheel about a gyroscopewheel axis that is non-parallel to the cage axis, the gyroscope wheeldisposed within the cage; producing a moment on the aircraft uponrotating the gyroscope wheel, wherein the cage and gyroscope wheelpartially form a first vibration control assembly; and controlling themoment produced by controlling an angular orientation of the gyroscopewheel.
 13. The method of claim 12, wherein the angular orientation ofthe gyroscope wheel is controlled by translating an arm operativelycoupled to the gyroscope wheel along a track formed along an innersurface of a structure that partially surrounds the gyroscope wheel. 14.The method of claim 13, wherein the structure and the gyroscope wheelare rotated synchronously.
 15. The method of claim 13, wherein therotational speed of the structure is controlled with a braking mechanismthat applies a braking force to the structure.
 16. The method of claim15, wherein a power source is provided power with energy generated bythe braking force applied to the structure.
 17. The method of claim 12,further comprising controlling the moment produced on the aircraft byvarying a rotational speed of the gyroscope wheel.